Bypass engines, such as jet turbine engines for aircraft, generally include a bleed air relief system for venting heated air from the engine compressor to the bypass air duct of the engine. Bleed air relief generally is provided to prevent the compressor from exceeding its surge limits during operation. Excess compressed air is bled off from the compressor upstream of the point at which the compressed air is mixed with fuel in the combustor section of the engine. Such bleed air relief systems typically include a bleed relief valve positioned along an air duct for controlling the release of bleed air from the compressor into the bypass air stream. The bleed air coming from the compressor generally is at a high pressure and temperature, typically upwards of approximately 950° F.-1,300° F., and often routed from the bypass air duct for deicing, pressurizing the cabin of an aircraft, operation of pneumatic actuators, and a variety of other applications.
The high temperature of the bleed air from the compressor can, when injected into the bypass air stream, cause damage to the walls of the bypass duct, translating sleeve, and cascades of the thrust reversers, and translating cowl portions of the engine, especially under conditions of prolonged exposure during thrust reverse operations of the engine. It is now known to utilize a perforated cover at the outlet of the bleed relief duct along the bypass duct of the engine to induce mixing of the bleed air with the bypass air for cooling of the bleed air prior to it impinging upon the walls of the bypass duct, translating sleeve, cascades, and other components of the engine. However, the heat reduction effect provided by such conventional perforated covers can be somewhat limited under certain high pressure and/or high velocity flow conditions.
One condition in which the heat reduction effectiveness of conventional perforated covers is lessened is where the bleed valve is in a failed open state. In this condition, the bleed valve that normally regulates the amount of air bled from the compressor can become stuck in the “open” position, thereby continuously injecting high pressure bleed air at temperatures of 950° F.-1,300° F. or greater into the bypass duct. The prolonged effect of high temperature bleed air can have a undesirable effect upon the several components encountered. Such temperatures can be higher than the curing temperatures of adhesives used in the construction of components such as the cascade translating sleeve, thereby leading to potential cracks and even failures. Similarly, in a forward mode operation of the aircraft, exposure of the translating cowl or inner walls of the bypass duct and other portions of the engine to such high temperature bleed air for several hours during flight can cause some damage or discoloration, such as scorching, cracks, delamination, dis-bonding and eventually failures of such components. If not corrected in a timely manner, eventual repair or replacement of engine components, e.g., translating cowl, may be required.
Accordingly, it can be seen that a need exists for a bleed air relief system that addresses the foregoing and other related and unrelated problems in the art.